XFOIL Version 6.96 Calculated polar for: NACA 63-209 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.200 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -9.000 -0.5555 0.08559 0.08214 -0.0190 1.0000 0.0479 -8.750 -0.5607 0.08025 0.07685 -0.0232 1.0000 0.0489 -8.500 -0.5725 0.07376 0.07040 -0.0298 1.0000 0.0493 -8.250 -0.5856 0.06859 0.06519 -0.0341 1.0000 0.0496 -8.000 -0.5983 0.06376 0.06006 -0.0383 1.0000 0.0517 -7.750 -0.6032 0.06183 0.05763 -0.0393 1.0000 0.0527 -7.500 -0.5543 0.04238 0.03883 -0.0396 1.0000 0.0558 -7.250 -0.5468 0.03883 0.03522 -0.0396 1.0000 0.0581 -7.000 -0.5409 0.03495 0.03113 -0.0397 1.0000 0.0619 -6.750 -0.5398 0.03003 0.02571 -0.0398 1.0000 0.0676 -6.500 -0.5258 0.02711 0.02288 -0.0392 1.0000 0.0705 -6.250 -0.5178 0.02432 0.01967 -0.0383 1.0000 0.0812 -6.000 -0.5044 0.02220 0.01767 -0.0372 1.0000 0.0866 -5.750 -0.5023 0.02905 0.02292 -0.0352 1.0000 0.0505 -5.500 -0.4807 0.02634 0.01949 -0.0326 1.0000 0.0405 -5.250 -0.4630 0.02275 0.01564 -0.0312 1.0000 0.0392 -5.000 -0.4440 0.02144 0.01407 -0.0295 1.0000 0.0399 -4.750 -0.4252 0.01921 0.01162 -0.0282 1.0000 0.0417 -4.500 -0.4057 0.01765 0.00998 -0.0268 1.0000 0.0428 -4.250 -0.3863 0.01654 0.00883 -0.0255 1.0000 0.0444 -4.000 -0.3665 0.01566 0.00791 -0.0243 1.0000 0.0466 -3.750 -0.3311 0.01493 0.00712 -0.0261 0.9962 0.0519 -3.500 -0.2933 0.01367 0.00591 -0.0286 0.9911 0.0590 -3.250 -0.2532 0.01281 0.00504 -0.0315 0.9854 0.0741 -3.000 -0.2196 0.01006 0.00414 -0.0347 0.9809 0.4679 -2.750 -0.1851 0.00965 0.00429 -0.0359 0.9737 0.6342 -2.500 -0.1453 0.00964 0.00435 -0.0380 0.9685 0.6899 -2.250 -0.1102 0.00967 0.00440 -0.0391 0.9606 0.7304 -2.000 -0.0741 0.00982 0.00465 -0.0399 0.9555 0.7760 -1.750 -0.0440 0.00989 0.00473 -0.0397 0.9462 0.8022 -1.500 -0.0094 0.00987 0.00468 -0.0406 0.9394 0.8177 -1.250 0.0208 0.00983 0.00461 -0.0407 0.9300 0.8315 -1.000 0.0495 0.00980 0.00456 -0.0406 0.9203 0.8444 -0.750 0.0780 0.00975 0.00449 -0.0402 0.9119 0.8563 -0.500 0.1024 0.00972 0.00445 -0.0390 0.9009 0.8687 -0.250 0.1261 0.00969 0.00441 -0.0377 0.8900 0.8816 0.000 0.1498 0.00965 0.00436 -0.0364 0.8799 0.8948 0.250 0.1735 0.00959 0.00429 -0.0350 0.8704 0.9085 0.500 0.1967 0.00954 0.00425 -0.0337 0.8591 0.9232 0.750 0.2225 0.00948 0.00419 -0.0329 0.8486 0.9381 1.000 0.2533 0.00942 0.00412 -0.0332 0.8394 0.9521 1.250 0.2895 0.00936 0.00407 -0.0347 0.8292 0.9646 1.500 0.3297 0.00932 0.00406 -0.0373 0.8184 0.9758 1.750 0.3715 0.00929 0.00404 -0.0402 0.8076 0.9864 2.000 0.4142 0.00924 0.00401 -0.0433 0.7959 0.9971 2.250 0.4370 0.00925 0.00404 -0.0426 0.7821 1.0000 2.500 0.4556 0.00929 0.00407 -0.0409 0.7664 1.0000 2.750 0.4769 0.00927 0.00402 -0.0394 0.7457 1.0000 3.000 0.4995 0.00921 0.00388 -0.0379 0.7164 1.0000 3.250 0.5234 0.00919 0.00380 -0.0368 0.6818 1.0000 3.500 0.5489 0.00928 0.00385 -0.0362 0.6530 1.0000 3.750 0.5742 0.00940 0.00392 -0.0354 0.6162 1.0000 4.000 0.5978 0.00964 0.00398 -0.0344 0.5512 1.0000 4.250 0.6142 0.01068 0.00418 -0.0324 0.3629 1.0000 4.500 0.6225 0.01368 0.00551 -0.0305 0.0805 1.0000 4.750 0.6454 0.01467 0.00644 -0.0299 0.0596 1.0000 5.000 0.6679 0.01575 0.00754 -0.0291 0.0515 1.0000 5.250 0.6905 0.01681 0.00860 -0.0283 0.0466 1.0000 5.500 0.7115 0.01866 0.01043 -0.0273 0.0434 1.0000 5.750 0.7370 0.01950 0.01138 -0.0268 0.0403 1.0000 6.000 0.7623 0.02086 0.01282 -0.0263 0.0383 1.0000 6.250 0.7881 0.02250 0.01457 -0.0257 0.0371 1.0000 6.500 0.8141 0.02447 0.01672 -0.0252 0.0364 1.0000 6.750 0.8396 0.02689 0.01944 -0.0244 0.0365 1.0000 7.000 0.8630 0.02887 0.02160 -0.0240 0.0349 1.0000 7.250 0.8852 0.03162 0.02466 -0.0231 0.0345 1.0000 7.500 0.9004 0.03828 0.03221 -0.0206 0.0415 1.0000 10.250 0.7939 0.08696 0.08383 -0.0103 0.0557 1.0000 10.500 0.7738 0.09312 0.09006 -0.0131 0.0558 1.0000 10.750 0.7510 0.09998 0.09698 -0.0172 0.0558 1.0000